Apparatus and methods for in-space satellite operations

ABSTRACT

Apparatus and methods for performing satellite proximity operations such as inspection, recovery and life extension of a target satellite through operation of a “Satellite Inspection Recovery and Extension” (“SIRE”) spacecraft which can be operated in the following modes (teleoperated, automatic, and autonomous). The SIRE concept further consists of those methods and techniques used to perform certain (on-orbit) operations including, but not limited to, the inspection, servicing, recovery, and lifetime extension of satellites, spacecraft, space systems, space platforms, and other vehicles and objects in space, collectively defined as “target satellites”. The three basic types of SIRE proximity missions are defined as “Lifetime Extension”, “Recovery”, and “Utility”. A remote cockpit system is provided to permit human control of the SIRE spacecraft during proximity operations.

RELATED APPLICATIONS

This application is a continuation of my pending U.S. application Ser.No. 09/489,140, filed Jan. 21, 2000, which is a continuation of my U.S.application Ser. No. 09/146,161, filed Aug. 2, 1998, now U.S. Pat. No.6,017,000, which is a divisional application of U.S. application Ser.No. 08/700,712, filed Jul. 12, 1996, now U.S. Pat. No. 5,806,802, whichis a national phase application of PCT International Application NumberPCT/US95/15103, filed Nov. 13, 1995, which is, in-turn, acontinuation-in part of PCT International Application NumberPCT/US94/13052, filed Nov. 14, 1994, which is, in-turn, acontinuation-in-part of my U.S. application Ser. No. 08/152,459, filedNov. 12, 1993, now U.S. Pat. No. 5,511,748.

BACKGROUND OF THE INVENTION

This invention pertains to apparatus and methods for in-space satelliteoperations, such as modifying the useful life of a space satellite,inspecting it, adjusting its space trajectory, and the like.

More particularly, the invention relates to such apparatus and methodsfor extending or otherwise modifying the useful operational lifetime ofsatellites which perform communications, weather reconnaissance, spacereconnaissance and similar functions.

In another respect, the invention pertains to such apparatus and methodsfor extending the useful life of such satellites without performingcomplicated in-space refueling or repair functions.

According to still another aspect, the invention pertains to apparatusand methods for effecting planned deorbit and reentry of a satellite orreboosting a spent satellite to a higher orbit or another trajectory, todelay deorbit or to place a satellite in a parking or other operationalor non-operational orbit or trajectory. Similarly, the invention relatesto effecting the deorbit or altering the trajectory of debris in spaceto avoid collisions with other spacecraft.

In yet another aspect, the invention pertains to apparatus and methodsfor performing a variety of proximity operations, e.g., inspection of anoperational or non-operational satellite, to determine its status, etc.

In still another aspect, the invention pertains to apparatus and methodsfor delivering or replenishing supplies to orbiting spacecraft such asthe planned International Space Station.

Because of the high reliability of contemporary electronics, theend-of-life (EOL) of most satellites is caused by on board propellantdepletion and the corresponding loss of attitude and position control,i.e., for orientation, pointing, including stabilization, and orbitcontrol. The previous proposed approach to extending EOL is to replenishthe propellant in the satellite tanks by refueling from anotherspacecraft. Alternatively, mechanical attachment of additional externalpropellant tanks to the target satellite would also accomplish thisobjective.

In addition to EOL by normal propellant depletion, there have beennumerous instances in which satellites have been initially delivered tounacceptable orbits. These orbits have been corrected by additionalpropulsion maneuvers. However, use of the satellites' onboard propellantto move it to an acceptable orbit resulted in a corresponding reductionin the useful life of the satellite. In some instances, initial orbitcorrection was impossible because it would have completely depleted thesatellite's onboard propellant supply.

In the past, considerable effort has been expended to develop in-spacerefueling technology. However, this has required extensive and expensivemodifications to conventional satellites, risky proximity operations,possible contamination of the satellite by escaping fuel and otherpractical problems.

Conversely, when extension of the operating life of a satellite cannotbe effected for various reasons, e.g., other malfunctions of a satelliteor its equipment which cannot be repaired or obsolescence of thesatellite, it would be desirable to be able to effect pre-planneddeorbit and reentry. In this way, the inoperable or obsolescentsatellite will not continue to clutter the available space for workingsatellites and reduce the likelihood of collision with other satellitesor space vehicles. If the deorbit and reentry can be preplanned, thesetechniques will also reduce the possibility of reentry into populatedareas with possible disastrous results. Furthermore, even if planneddeorbit and reentry is not necessary or if it would be desirable tootherwise change the space trajectory of a satellite, it would bedesirable to provide apparatus and methods for changing the spacetrajectory of a satellite to another operational or non-operationaltrajectory or for reboosting satellites which are cluttering usefulorbits or which are about to deorbit, into less cluttered and lessdangerous parking orbits.

SUMMARY OF THE INVENTION

The principal object of the present invention is to provide apparatusand methods for in-space satellite operations, such as, for example,extending or otherwise modifying the useful life of a space satellite,modifying its space trajectory, etc.

Yet another object of the invention is to provide such extension of theuseful life of a space satellite by a simplified method and usingsimplified apparatus in comparison to prior art techniques which involverefueling the space satellite.

Still another object of the invention is to provide apparatus andmethods which permit planned deorbit and reentry of spent or obsoletesatellites, which permit changing the space trajectory of a satellite toanother operational or non-operational trajectory or which permitreboosting spent or obsolete satellites to a parking orbit.

These, other and further objects and advantages of the invention will beapparent to those skilled in the art from the following detaileddescription thereof, taken in conjunction with the drawings.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a perspective view of an extension spacecraft, configured inaccordance with a presently preferred embodiment of the invention;

FIG. 2 is a partially cut-away perspective view of the service module ofthe extension spacecraft of FIG. 1;

FIG. 3 is a perspective view of the command module of the extensionspacecraft of FIG. 1;

FIG. 4 illustrates the docking maneuvers and mechanical interconnectionof the extension spacecraft of FIGS. 1-3 with a target satellite;

FIGS. 5-7 illustrate a typical mission scenario performed by theapparatus and method of the invention, to transfer a satellite from anunusable orbit to its intended operational orbit and thereafter providestationkeeping and pointing for the docked combination extensionspacecraft-target satellite; and

FIGS. 8-9 illustrate a remote cockpit system employed in the preferredembodiment of the invention to provide human control of proximityoperations such as docking, inspection, etc.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS

As used herein the term “adjusting the life of a target satellite” meanseither extending the useful life of a target satellite, which isnormally limited by the availability of onboard propellant for positioncontrol, or terminating the orbiting activity of a spent or obsoletesatellite by planned deorbit and reentry or by transferring a satellitefrom its previous orbit another trajectory or to a parking orbit.

As used herein the term “controlling the position of the dockedsatellite-spacecraft combination” includes both controlling thetrajectory of the docked combination relative to the earth and/orcontrolling the attitude of the docked combination relative to the earthor to the star field.

Briefly, in accordance with one embodiment of the invention, I provideapparatus and methods for performing satellite proximity operations suchas inspection, recovery and life extension of a target satellite throughoperation of a “Satellite Inspection Recovery and Extension” (“SIRE”)spacecraft which can be operated in the following modes (teleoperated,automatic, and autonomous). The SIRE concept further consists of thosemethods and techniques used to perform certain (on-orbit) operationsincluding, but not limited to, the inspection, servicing, recovery, andlifetime extension of satellites, spacecraft, space systems, spaceplatforms, and other vehicles and objects in space, collectively definedas “target satellites”.

The three basic types of SIRE proximity missions are defined as“Lifetime Extension”, “Recovery”, and “Utility”. Each type of mission isfurther separated into additional categories depending on more specifictechnical and operational requirements. For example, the objective ofthe Lifetime Extension Mission is to provide additional stationkeepingpropellant for satellites that are approaching their projected end life(EOL) due to onboard propellant depletion but which are otherwise fullyfunctional. The Lifetime Extension Mission thus enables the fullyfunctional satellite to remain operational in its desired (revenueproducing) orbit for an extended period beyond its projected end of lifeby forming a docked SIRE satellite-spacecraft combination.

To perform the Life Extension Mission, the SIRE spacecraft includesguidance, navigation and control systems, an onboard propellant supplyand docking means for mechanically connecting the target satellite andthe SIRE spacecraft to form the docked satellite-spacecraft combination.Preferably, the propulsion system is hypergolic consisting ofmono-methylhydrazine and N₂O₄ for proximity operations. The guidance,navigation and control systems of the SIRE spacecraft provide the meansfor controlling the position of the docked satellite-spacecraftcombination. The onboard propellant supply is sufficient to provide forrendezvous and docking of the SIRE spacecraft with the target satelliteand for position control of the docked satellite-spacecraft combination.

In accordance with another embodiment of the invention, I provide amethod for adjusting the life of a target satellite. The method of theinvention comprises the steps of mechanically connecting a SIREspacecraft to the target satellite, forming a dockedsatellite-spacecraft combination and activating the guidance, navigationand control systems of the SIRE spacecraft to provide position controlfor the docked satellite-spacecraft combination. The SIRE spacecraftused in this method includes onboard propellant supply for positioncontrol of the docked satellite-spacecraft, after docking. By having theSIRE spacecraft perform all stationkeeping functions such as positionand attitude control, the satellite may perform its designed functions,such as telecommunications and weather mapping, long after its originalprojected end of life.

The objective of the “Recovery Mission” is to correct various anomaliesencountered by orbiting satellites. These anomalies include incorrectlaunch orbit, orbital decay, loss of satellite function capability, andsatellite system failure. To correct incorrect launch orbit or orbitaldecay, the SIRE spacecraft is similar in construction to the spacecraftutilized for Lifetime Extension Missions. The “recovery” SIRE spacecraftincludes guidance, navigation and control systems, an onboard propellantsupply and docking means for mechanically connecting the targetsatellite and the SIRE spacecraft to form the dockedsatellite-spacecraft combination. The guidance, navigation and controlsystems of the SIRE spacecraft provide the means for controlling theposition of the docked satellite-spacecraft combination. The onboardpropellant supply is sufficient to provide for rendezvous and docking ofthe spacecraft with the satellite and for subsequently transferring thesatellite to another orbit or effecting reentry of thespacecraft-satellite combination.

Where the target satellite has suffered a loss of capability such asnon-deployed antenna or solar array, the SIRE spacecraft includesadditional apparatus for effecting repair of refurbishment, or foreffecting return of the spacecraft to Earth (Shuttle). In a preferredembodiment, the SIRE spacecraft includes docking means for effecting thedocked spacecraft-satellite combination and intervention tools foreffecting repair and/or refurbishment of the target satellite. Theintervention tools, by way of example, include means for removing andreplacing sections of the satellite such as spacecraft thermal blankets,means for severing restraint cables that prevent deployment of theantenna or solar array and means for deploying stuck mechanisms. Moreparticularly, by way of example these intervention tools may include, asatellite capture bar such as the “Stinger” designed by NASA, one ormore robotic arms similar to a smaller version of the Remote ManipulatorSystem (RMS) found on the Space Shuttle, a satellite closeup inspectiontool such as a remotely operated camera, a two-finger gripper, acable/pin/bolt cutter and a lever force tool.

Alternatively, for certain system (subsystem, component) failures, theRecovery Mission provides the target satellite with a substitute orsupplemental system to “recover” the satellite to it designatedoperational capability including redundancy. For example, Anik E-1 andAnik E-2, which are three-axis stabilized spacecraft designed to providetelevision coverage for Western Europe, have both suffered failures oftheir primary momentum wheels. Similarly, several additional spacecrafthave also encountered anomalies with their momentum wheels indicatingpossible premature failures. The docking of a SIRE spacecraft equippedwith supplemental momentum wheels to the target satellite would providethe necessary stabilization to enable the satellite to remainoperational for its projected end of life.

The objective of the “Utility Missions” include all other SIRE in-spacemaintenance missions such as in-space inspection of satellites and otherspace vehicles and objects, environmental protection and collisionavoidance (debris control), and resupply of orbiting satellites such asthe planned International Space Station Alpha (ISSA). An additionalUtility Mission would entail the planned “decommissioning” of anorbiting spacecraft. Under certain circumstances, such as militaryapplications, it would be desirable to deactivate a satellite. In apreferred embodiment, the SIRE spacecraft includes a liquid contaminantthat is ejected through nozzles or the like from the SIRE spacecraft tocontact the target satellite. Ideally, the contaminant degrades systemperformance of the target satellite to eventually effect completefailure of the target satellite. A unique feature of the SIRE spacecraftis the hypergolic propulsion system used for rendezvous. The spraying ofthis propellant (mono methylhydrazine and N₂O₄) on a target satellitewould cause contamination of the spacecraft systems, e.g. rapiddeterioration of the solar cells or telecommunication disks, to effectthe degradation and eventual failure of the target satellite.

Turning now to the drawings, FIGS. 1-3 illustrate a SIRE spacecraftconstructed in accordance with and used in accordance with theprinciples of the present invention. The spacecraft 10 comprises acommand module 11 and a service module 12. The SIRE satellite embodiesexoatmospheric construction and is adapted to be carried into space,e.g., to a rendezvous phasing orbit or low Earth orbit in the enclosedcargo bay or within the enclosing shroud of an earth launch vehicle(ELV) such as, for example, the Taurus or the Space Shuttle, dependingon mission requirements, availability, cost, etc. For example, in oneembodiment of the invention, the baseline earth launch vehicle is theDelta 7920, which has a low earth orbit payload insertion capability ofapproximately 5,000 kilograms and a geosychronous transfer orbitcapability of approximately 1,300 kilograms.

The service module 12 operates as a “space bus” for the command module11, providing among other functions, propulsion, power andcommunications support, thus minimizing the requirements forcorresponding subsystems in the command module 11. The operations phasedesign lifetime of the command module 11 for in-space servicing cantherefore be relatively short, based on specific programmed tasks at thetarget vehicle during a fixed period of activity. For certain missions,the command module 11 will separate from the service module 12 andoperate independently. Also, for certain missions, a space transfervehicle (STV), such as that disclosed in my issued U.S. Pat. No.5,242,135, can be employed to transfer the SIRE spacecraft 10 from thelaunch insertion orbit to a rendezvous phasing orbit (RPO).

As will be apparent to those skilled in the art, all of the functions ofthe command module 11 could be incorporated into the service module 12,although the separate command and service modules herein describedprovide for maximum mission flexibility and are, accordingly, apresently preferred embodiment of the invention.

Referring more particularly to FIG. 2, the primary purpose of theservice module 12 is to augment the propulsion capabilities of thecommand module 11. For example, if the command module 11 is configuredas a variant of the SDIO Lightweight Exoatmospheric Projectile (LEAP)Vehicle, the service module 12 can be based on the design of theexisting “Small Altimeter” (SALT) satellite manufactured for the UnitedStates Navy by Intraspace, Inc., North Salt Lake City, Utah. The servicemodule 12 includes a command module adapter ring 21, GPS antenna 22,S-Band OMNI antenna 23, orbit insertion motors 24, propellant tanks 25,batteries, 26. Mounted on the mid-deck 27 is a reaction control system28 and on-board processor 29. These components are enclosed by amonocoque structure 30, on which are mounted solar power cell arrays 31.

The service module 12 is sized to perform all rendezvous and proximitymaneuvers, as well as specific transfer maneuvers required for the SIREspacecraft-target satellite docked combination. For certain targetspacecraft locations, the energy requirements to position the SIREspacecraft for rendezvous may be greater than that available from theservice module 12, for example, an inclination change for the targetsatellite. In such cases, the STV would be added to the SIRE spacecraft10 to augment the propulsion capabilities of the service module 12.

For major maneuvers, the service module 12 is equipped with a storablebiopropellant system consisting of a “quad” array of four upratedMarquardt R-4-D 490 Newton (100 lb.) thrust axial engines. Thisconfiguration provides adequate thrust-to-weight ratio to minimize theeffects of non-impulsive maneuvers, as well as redundancy for engine-outcapability to complete the mission. Marquardt R-4-D engines are selectedfor their very high reliability, high Isp (322 seconds), maturity (over800 produced) and availability.

To prevent contamination of the target satellite when the SIREspacecraft is stationkeeping, the SIRE spacecraft attitude controlsystem is a nitrogen cold gas system consisting of 16×5 lb. thrustersmounted in quads on the circumference of the service module 12. Thisconfiguration enables both three-axis rotation and three-axistranslation for example, for stationkeeping and docking.

Referring more specifically to FIG. 3, the command module 11 includesseveral major subsystems, including guidance, navigation and control(GNC) system used for all SIRE spacecraft operations, a main propulsionsystem with “divert” thrusters of approximately 100 lbs. (490 N) thrusteach, an attitude control system, and data and communication subsystems.The command module payload consists of a “seeker” subsystem with sensorsfor target location, tracking and inspection, and a docking system withvarious servicing devices such as a docking apparatus or robotic armswith clamps or grippers.

The basic configuration of the command module 11 is defined as acompletely independent vehicle to enhance mission planning flexibility,minimize interface requirements, maximize the use of existing ordevelopmental small spacecraft, and enable independent testing andverification of certain proximity operations and hardware in groundfacilities prior to launch. The command module 11 may remain attached tothe service module 12 (as for the UHF-1 recovery mission, describedbelow), or it may be detached to operate autonomously. The servicemodule 12 could, therefore, carry two or more command modules 11. Insuch configuration, the service module 12 acts as the primary spacecraftand the command module or modules can be detached fore use asobservation spacecraft. In either case, prior to separation of thecommand module(s) 11, certain rendezvous braking maneuvers would beperformed by the divert thrusters of the combined command module-servicemodule.

The baseline design command module 11 consists of a variant of the SDIOLEAP with minor modifications. The Rocketdyne AHIT Vehicle is selectedas the baseline command module 11. This vehicle has completed severalfull-up hover tests in the SDIO National Hover Test Facility. In currentconfiguration it weighs 10.2 kilograms, including 1.7 kilograms ofpropellant. It produces a delta velocity increment of 357 m/sec.

In this configuration, the command module includes cold gas attitudecontrol system thrusters 32 and two divert thrusters 33 which havesignificantly higher thrust (490 N, 100 lb.) than the service moduleengines (5 lb.). These divert thrusters 33 are aligned along the line ofsight from the service module 12 toward the target satellite. Thesedivert thrusters 33 would not be used in close proximity to the targetsatellite to preclude contamination of the satellite. The remaining twodivert thrusters of the AHIT vehicle are removed.

This forward alignment of the divert thrusters enables the seekerassembly to be continuously oriented toward the target satellite, thusprecluding the necessity of rotating the SIRE spacecraft 180 degreesopposite to the target line of site to perform braking maneuvers.Although the engines 24 of the service module 12 could be used toperform braking, the low thrust level of these engines (20 lbs. total)would result in much longer burn times and very narrow margins inignition time, burn durations, orbital position, and relative velocity.

FIGS. 5-7 illustrate a typical mission scenario which can beaccomplished by the apparatus and methods of the present invention.Illustratively, this scenario envisions the recovery of the Navy UHF-1satellite which was launched into a non-operational orbit on Mar. 29,1993, by a degraded launch vehicle. Subsequently, the Navy stated thatthe UHF-1 satellite is a total loss. At present, the UHF-1 satellite 41is in essentially a geosynchronous transfer orbit 51 with a perigee at118 nm, apogee at 19,365 nm and an inclination at 27 degrees. Therecovery flight profile depicted in FIGS. 5-7 is designed to accomplishinsertion of the satellite 41 into geostationary orbit (GEO) 52 bycircularizing the orbit and reducing its inclination to approximatelyzero degrees.

To accomplish this mission, the SIRE spacecraft 10 is launched from theEarth by an earth launch vehicle 53, into a Rendezvous Phasing Orbit(RPO) 54 with a perigee of 180 nm, an apogee of approximately 19,345 nmand an inclination of 27 degrees. After insertion of the SIRE spacecraft10 into RPO, a four-impulse sequence is initiated which consists ofcoeliptic sequence initiation (CSI), constant delta height (CDH),terminal phase initiation (TPI) and braking. CSI establishes a desiredratio of relative height to phase angle between the SIRE spacecraft 10and the target satellite 41. CSI also establishes, based on subsequentmaneuvers, the standard lighting conditions as well as transfer time forthe final approach to the target 41. CDH establishes a constantdifferential altitude between the SIRE spacecraft 10 and the targetsatellite. TPI establishes a SIRE spacecraft trajectory that willintercept the target satellite 41 at a specific time and position on theorbit 52 of the target satellite 41. A nominal transfer interval of 130degrees is used to optimize propellant usage, provide adequate controlauthority during the final approach, insure the apparent inertial motionof the target satellite 41 (relative to the star field) as near zeroduring the latter part of the intercept, and insure that the transfer isalong the line of sight. Braking is performed as a series of distinctmaneuvers performed at specific range/rate “gates”, each of which occursat a range from the target where the actual range/rate is reduced to apreplanned value. The maneuvers at these gates gradually reduce therelative velocity between the vehicles to zero. After docking of theSIRE spacecraft 10 with the target satellite 41, the docked combination57 then perform a series of maneuvers to raise the perigee of the dockedcombination 58 through intermediate orbits (indicated by the dash lineson FIG. 7) to raise the perigee to 19,365 nm and reduce the inclinationto near zero, placing the docked combination in final operational orbit(GEO) 52.

Of particular importance, in a preferred embodiment of the invention,the method and apparatus include use of a “remote cockpit” whichcontains a guidance computer, hand controllers and a visual display. Aswould be understood by those in the art, the remote cockpit may belocated anywhere such as the Space Shuttle, Space Station, MissionControl Center, mountain bunker, or even mobile van or ship to provide“remote” control of the SIRE spacecraft. In contrast to a ground stationcontrol center typically employed to control a satellite, spacecraft orthe like, the remote cockpit is configured substantially similar to thecockpit of an aircraft or spacecraft. In order to overcome the problemsof ground station control of a spacecraft which does not provide theintimate control beneficial for a spacecraft performing proximityoperations, I provide a remote cockpit including an image monitor, anFDAI (graduated 3-axis “8 ball”, analog or digital readouts of downlinkinputs from the SIRE, a Rotation Hand Controller (“RHC”) and aTranslation Hand Controller (“THC”). The image monitor provides apilot's eye window view from the SIRE spacecraft. The FDAI, similar tothe system employed by the Gemini and Apollo programs, provides acentralized inside-out display of the spacecraft attitude, vehiclerates, and attitude errors. More particularly, the FDAI providesattitude (roll—360 degrees, pitch—360 degrees, and yaw—+/−80 degrees)relative to an inertial frame of reference centered on the objectsatellite at the nominal time the rendezvous braking is completed.Angular rates about each of the three mutually perpendicular spacecraftbody axes, role, pitch and yaw are provided to the pilot. Further, theFDAI provides attitude errors of the spacecraft's angular positionrelative to the inertial reference in all three axes.

To permit the “pilot” to remotely fly the SIRE, the remote cockpitincludes analog and digital readouts disclosing: range to the targetsatellite, range rate to the target satellite, line of sight (LOS) angleof the SIRE spacecraft relative to the target satellite, elapsed time,propellant remaining, and velocity (Delta V) imparted to the SIREspacecraft as a result of thruster commands along each of the three SIREbody axes.

To control the SIRE spacecraft, the remote cockpit preferably includes aRotation Hand Controller and a Translation Hand Controller. The RotationHand Controller provides three-axis manual body-rate commands for thespacecraft's rotational motion in both directions about the threeprincipal axis proportional to hand controller deflection. The RightHand Controller is capable of simultaneous multiaxis actuation whileattitude hold is established when the Right Hand Controller is placed inneutral. Preferably, the Right Hand Controller also provides emergencyangular acceleration commands directly to the RCS jet solenoids by“direct switches” at the limits of the hand controller movement (stops).

The Translation Hand Controller provides manual on/off accelerationcontrol to the thrusters due to on/off response from the TranslationHand Controller along each left/right, up/down, fore/aft hand direction.Accordingly, movement of the Translation Hand Controller results in SIREspacecraft translation (rectilinear motion) in both directions alongeach of the principal body axes (X, Y and Z).

Uplink signals from the guidance computer, based on the “control laws”and inputs from the human-operated hand controllers (in response tovisual displays generated from spacecraft onboard television cameras andother processed downlink inputs, such as attitude, consumables,caution/warning, etc.) provide teleoperation of the spacecraft duringproximity operations, e.g., docking, inspection, etc.

FIGS. 8-9 depict the system and methods of the presently preferredembodiment of the invention, which includes a “remote cockpit” 60 forhuman control of rendezvous and proximity operations, e.g., inspectionand/or docking. The remote cockpit 60, which can be located on theground or on board another spacecraft, includes the Guidance, Navigationand Control System computer 61 containing the control laws, a controlssystem 67 and variables display 62. The controls system 67 includespilot operated translation hand controller 63, rotational handcontroller 64, and engine on-off switches 65. Appropriate signalprocessors 66 are provided to generate the uplinked control signals 68and downlinked spacecraft motion variables signals 69. Processed controlsignals 71 (“pilot input”) from the controls system 67 and itsassociated signal processor 66, are uplinked to the spacecraft 72,which, in-turn, provides downlinked motion variables signals 69 to theGNC computer 61 to generate the display variables signals 73 to thesignal processor 66 of the displays system 62.

In a preferred embodiment, the present invention includes additionalapparatus and methods for closed-loop telepresence, or image processingwherein the movement of the extension spacecraft relative to the targetsatellite is controlled by uplink signals from human operated handcontrollers in a remote cockpit (which includes the guidance computer),in response to visual or image displays generated from spacecraftonboard cameras or the like, to provide teleoperation of the SIREspacecraft during proximity operations, such as docking, inspection,etc.

To control the SIRE spacecraft in proximity operations, the humanoperator is provided with a visual image similar to the view from awindow or portal of an aircraft or spacecraft. This image is produced byimage signals relayed to the remote cockpit from sensors located on theSIRE spacecraft. In a preferred embodiment, the sensors are one or morevideo cameras which relay video images to the remote cockpit. Additionalimage producing systems include the use of radar, laser imaging or thelike to create a three dimensional image useful for the “pilot” of theSIRE spacecraft for conducting proximity operations.

In a preferred embodiment, the visual image system shows the actual starfield background of at least fourth magnitude stars during twilight anddarkness in orbit. Providing a visible star field is very important tothe pilot's rendezvous capabilities. As a result of the geometryestablished during the initial part of the rendezvous, the finalvelocity match of the SIRE spacecraft with the target satellite isachieved by maneuvering directly along the line-of-sight of the SIRE tothe target satellite. Proper lighting results in darkness during whichnay apparent relative motion of the target against the background starfield is nulled by vertical or horizontal translation of the SIRE. Therendezvous geometry enables a straightforward visual/manual thrustingprofile using relatively low thrust attitude control thrusters. Theprimary parameter throughout the rendezvous, particularly during theterminal phase, is the relative motion between the active SIREspacecraft and the passive target satellite wherein the target satelliteappears fixed in space.

When the relative motion of the target satellite against the star fieldis zero (up/down and left/right), the pilot is on the proper intercepttrajectory with the target. As an example, is the target moves uprelative to the star field, the pilot will thrust “up” with theTranslation Hand Controller to null the relative motion between the SIREand the target satellite and gain on the intercept trajectory. Duringthis process, and using the FDAI, the pilot maintains a fixed spacecraftattitude with the Right Hand Controller. The pilot with his eyes andbrain processor is essential to these maneuvers, and accordingly, as isthe need for a window (visual image showing stars), and handcontrollers.

In an additional embodiment, the imaging system of the present inventionincludes a Geometrical Pattern Recognition System (“GPRS”). Inoperation, the GPRS functions substantially similar to the Apollo“COAS”/“stand-off cross” target system. For more detail regarding theApollo “COAS”/“stand-off cross” target system, refer to the ApolloExperience Report- “Crew Provisions and Equipment Subsystem” McAllister,F. A., NASA/MSC, March 1972, which is incorporated herein by reference.By means of the GPRS, when approaching the target satellite to performproximity operations, such as docking, the pilot in the remote cockpitis presented with two computer generated 3-dimensional perspectivegeometrical line patterns (figures) of the target satellite, the firstbeing a representation of the target's current (real-time) position, andthe second being a representation of the target satellite at the point(time) of initial contact with the SIRE spacecraft. The real-time targetrepresents the actual precise geometrical orientation of the targetsatellite relative to the pilot's eyes as the target appears to move tothe point of initial contact. The point (time) of contact targetrepresents the required precise geometrical orientation of the targetsatellite relative to the pilot's eyes at the time (point) of initialcontact between the two vehicles. In this manner, although the SIRE isactually moving relative to the target satellite, the frame of referenceis fixed at the pilot's eyes.

For the SIRE to successfully capture the target (dock), the two patternsmust be superimposed at the time/point of initial contact, or a zerorelative distance between the SIRE and the target satellite. Therelative position and orientation of the target is availablecontinuously from the vectors of the two vehicles.

In an additional embodiment, the GPRS provides a third image to thepilot's eyes, a “nominal target”. The nominal target is the nominal(planned) geometrical orientation of the target spacecraft as the targetmoves to the point of initial contact. Preferably, this image can betuned on and off to allow the pilot to superimpose the nominal locationand the orientation of the target satellite at any given time betweenthe beginning of the approach until SIRE and satellite engagement. Tofurther facilitate pattern recognition, each of the three images (point(time) of contact, real-time target and nominal target) is presented ina different color or represented by a different line type, e.g., shortdashes or dots.

Having described by invention in such terms as to enable those skilledin the art to understand and practice it, workers skilled in the artwill recognize that changes may be made in form and detail withoutdeparting from the spirit and scope of the invention,

I claim:
 1. A remote cockpit system for controlling in-space proximityoperations of a spacecraft in proximity to a target satellite,comprising: (a) a cockpit remote from said spacecraft; (b) a spacecraftvariables display system located in said remote cockpit, said spacecraftvariables display system including a visual display for signalsdownlinked from an image producing means and motion sensor means onboardsaid spacecraft; and (c) control means adapted to be manually operatedby a human pilot in said remote cockpit in response to displays on saiddisplay system, including means to generate pilot input signals throughsaid computer which provide control attitude and velocity uplinked tosaid spacecraft to control said spacecraft during said proximityoperations.
 2. The remote cockpit system of claim 1 wherein saidspacecraft variable display system further comprises; a geometricalpattern recognition system.
 3. The remote cockpit system of claim 2wherein said remote cockpit system further comprises; (a) a rotationhand controller; and (b) a translation hand controller.
 4. The remotecockpit of claim 1 further comprising a Flight Director AttitudeIndicator (FDAI) providing spacecraft roll, pitch and yaw.
 5. A methodof performing in-space proximity operations, said method comprising: (a)remotely controlling a Satellite Inspection Recovery and Extensionspacecraft in proximity to an orbiting target satellite, said SatelliteInspection Recovery and Extension spacecraft including, an interventiontool for refurbishing or repairing said satellite; and (b) operatingsaid intervention tool to refurbish or repair said satellite.
 6. Amethod of performing in-space proximity operations, said methodcomprising: (a) remotely controlling a Satellite Inspection Recovery andExtension spacecraft in proximity to an orbiting target satellite, saidSatellite Inspection Recovery and Extension spacecraft including, (i)guidance, navigation and control systems for position control of saidSatellite Inspection Recovery and Extension spacecraft, (ii) an onboardpropellant supply, (iii) deactivation means for deactivating a targetsatellite; and (b) deactivating said target satellite using saiddeactivation means of said Satellite Inspection Recovery and Extensionspacecraft.
 7. The method of performing in-space proximity operations ofclaim 6, wherein the step of deactivating said target satelliteincludes: said Satellite Inspection Recovery and Extension spacecraftejecting a solid, liquid or gaseous material upon said satellite.
 8. Amethod of performing in-space proximity operations, said methodcomprising: (a) remotely controlling a Satellite Inspection Recovery andExtension spacecraft in proximity to an orbiting target satellite, saidSatellite Inspection Recovery and Extension spacecraft including,inspection means for inspecting said target satellite; and (b)inspecting said target satellite using said inspection means of saidSatellite Inspection Recovery and Extension spacecraft.